This invention relates generally to improving the durability of gas turbine engine components and, particularly, in reducing the thermal stresses in the turbine engine stator components such as nozzle segments.
In a typical gas turbine engine, air is compressed in a compressor and mixed with fuel and ignited in a combustor for generating hot combustion gases. The gases flow downstream through a high pressure turbine (HPT) having one or more stages including one or more HPT turbine nozzles and rows of HPT rotor blades. The gases then flow to a low pressure turbine (LPT) which typically includes multi-stages with respective LPT turbine nozzles and LPT rotor blades. The HPT and LPT turbine nozzles include a plurality of circumferentially spaced apart stationary nozzle vanes located radially between outer and inner bands. Typically, each nozzle vane is a hollow airfoil through which cooling air is passed through. Cooling air for each vane can be fed through a single spoolie located radially outwardly of the outer band of the nozzle. In some vanes subjected to higher temperatures, such as the HPT vanes for example, an impingement baffle may be inserted in each hollow airfoil to supply cooling air to the airfoil.
The turbine rotor stage includes a plurality of circumferentially spaced apart rotor blades extending radially outwardly from a rotor disk which carries torque developed during operation. Turbine nozzles, located axially forward of a turbine rotor stage, are typically formed in arcuate segments. Each nozzle segment has two or more hollow vanes joined between an outer band segment and an inner band segment. Each nozzle segment and shroud hanger segment is typically supported at its radially outer end by flanges attached to an annular outer casing. Each vane has a cooled hollow airfoil disposed between radially inner and outer band panels which form the inner and outer bands. In some designs the airfoil, inner and outer band portions, flange portion, and intake duct are cast together such that the vane is a single casting. In some other designs, the vane airfoils are inserted in corresponding openings in the outer band and the inner band and brazed along interfaces to form the nozzle segment.
Certain two-stage turbines have a cantilevered second stage nozzle mounted and cantilevered from the outer band. There is little or no access between first and second stage rotor disks to secure the segment at the inner band. Typical second stage nozzles are configured with multiple airfoil or vane segments. Two vane designs, referred to as doublets, are a very common design. Three vane designs, referred to as Triplets are also used in some gas turbine engines. Doublets and Triplets offer performance advantages in reducing split-line leakage flow between vane segments. However, the longer chord length of the outer band and mounting structure compromises the durability of the multiple vane segment nozzles. The longer chord length causes an increase of chording stresses due to the temperature gradient through the band and increased non-uniformity of airfoil and band stresses, such as for example, shown in FIG. 6 for a conventional outer band. The increased thermal stress may reduce the durability of an outer band and the turbine vane segment. It is desirable to have a flange design for supporting turbine engine components such as the turbine nozzle segments that avoid reduction in the durability of multiple vane segments due to longer chord length of the outer band and mounting structure. It is also desirable to have turbine nozzle segments that avoid increase of chording stresses due to temperature gradient through the outer band and increased non-uniformity of airfoil stresses due to longer chord length of the multiple vane segments. It is also desirable to have turbine nozzle segments that avoid increase of stresses near the middle vane of a Triplet or other multiple vane segments which limits the life of the segment.